Rocket engine with cryogenic propellants

ABSTRACT

A cyrogenic-propellant rocket engine includes: at least a first tank for a first liquid propellant; a second tank for a second liquid propellant; a third tank for an inert fluid; an axisymmetrical nozzle including a combustion chamber, a device for injecting first and second liquid propellants into the combustion chamber, a nozzle throat, and a divergent section; and a heater device including at least one duct for conveying the inert fluid and arranged outside the nozzle in immediate proximity thereof, but without making contact therewith, to recover energy of thermal radiation emitted when the rocket engine is in operation and to heat the inert fluid.

FIELD OF THE INVENTION

The present invention relates to a cryogenic-propellant rocket enginehaving at least a first tank for a first liquid propellant, a secondtank for a second liquid propellant, a third tank for an inert fluid,and an axisymmetrical nozzle comprising a combustion chamber, a devicefor injecting first and second liquid propellants into the combustionchamber, a nozzle throat, and a divergent section.

PRIOR ART

U.S. Pat. No. 6,658,863 describes in particular a system for storing andsupplying inert gas such as helium under pressure on board arocket-engine launcher vehicle for the purpose of pressurizing one ormore propellant tanks in order to deliver the propellant continuously tothe rocket engine and in order to maintain the structural integrity ofthe tank. Such a known system implements a heat exchanger that may makeuse of the hot gases from the propulsion system of the launcher as asource of heat. As an alternative, proposals are also made for usingelectrical heating. Such solutions are not appropriate for all types ofrocket engine and divergent nozzle, and in particular they are poorlyadapted to recovering energy from rocket engines operating in anexpander type cycle or using divergent nozzles made of compositematerial.

In general, use is often made in industry of heat exchangers that heat afluid by making use of the power available from a solid part, whichpower is recovered either by conduction (contact between two parts) orelse by convection (contact between a hot fluid and a wall).

OBJECT AND DEFINITION OF THE INVENTION

The invention seeks to remedy the above-mentioned drawbacks of the priorart and to make it possible in optimized and simple manner to recoverthe heat available from a rocket engine for the purpose of heating aninert fluid for pressurizing one or more propellant tanks, and also toreduce the on-board weight in the launcher fitted with the rocketengine.

In accordance with the invention, these objects are achieved by acyrogenic-propellant rocket engine having at least a first tank for afirst liquid propellant, a second tank for a second liquid propellant, athird tank for an inert fluid, and an axisymmetrical nozzle comprising acombustion chamber, a device for injecting first and second liquidpropellants into the combustion chamber, a nozzle throat, and adivergent section, the rocket engine being characterized in that itfurther comprises a heater device including at least one duct forconveying said inert fluid and arranged outside the nozzle in theimmediate proximity thereof, but without making contact therewith, inorder to recover the energy of the thermal radiation emitted when therocket engine is in operation and in order to heat said inert fluid.

The radiant power emitted by a rocket-engine divergent nozzle sectionmade of composite material may be greater than 250 kilowatts per squaremeter (kW/m²) of the divergent section, and it is thus possible torecover a large amount of radiant energy for heating the inert fluid(e.g. helium available at 20 kelvins (K)), thereby pressurizing alauncher stage so as to obtain a direct increase in terms of payload (ofthe order of several tens of kilograms), without necessarily recoveringenergy from the rocket engine by conduction or convection.

According to an aspect of the invention, the heater device comprises ametal structure in which the inert fluid for heating flows and a finelayer that is strongly absorbent from a thermal radiation point of viewis deposited at least on the walls that face the nozzle constituting aradiant power source.

In a possible embodiment, the heater device comprises at least one torussurrounding the nozzle.

In yet another embodiment, the heater device comprises at least one tubewound helically around the divergent section of the nozzle.

In yet another advantageous embodiment, the heater device comprises twoto four tubes surrounding the divergent section of the nozzle.

Under such circumstances, the heater device may have four tubes, forexample.

Advantageously, the tubes are provided with means for providing acounterflow of the inert fluid in at least one tube.

The plurality of tubes may itself comprise a set of twisted-togethertubes wound around a common circular axis or a common helical axis.

The inert fluid is advantageously helium that may serve as apressurization gas, e.g. for pressurizing a liquid oxygen tank. Undersuch circumstances, connection pipes are arranged between the heaterdevice and the liquid oxygen tank in order to supply it withpressurizing helium.

The nozzle of the rocket engine may be made in a wide variety ofmanners, and may, for example, include an expandable divergent nozzlethat may advantageously be made of carbon-carbon composite material.

BRIEF DESCRIPTION OF THE DRAWINGS

Other characteristics and advantages of the invention appear from thefollowing description of particular embodiments given as examples andwith reference to the accompanying drawings, in which:

FIG. 1 is a diagrammatic cross-section view of a rocket engine nozzleand of a plate heater device in an embodiment of the invention;

FIG. 2 is an enlarged view of a portion II of the FIG. 1 plate heaterdevice;

FIG. 3 is a diagrammatic perspective view of the nozzle and of the FIG.1 plate heater device;

FIG. 4 is a diagrammatic perspective view of an embodiment of theinvention with a nozzle and a heater device of toroidal shape;

FIG. 5 is a diagrammatic perspective view of an embodiment of theinvention with a nozzle and a heater device in the form of a helicaltube;

FIG. 6 is a diagrammatic perspective view of an embodiment with a nozzleand a heater device having a plurality of helical tubes;

FIG. 7 is a diagrammatic perspective view of an embodiment of theinvention with a nozzle and a heater device having a plurality oftwisted tubes within a structure that can be inscribed in a torus; and

FIG. 8 is an overall diagrammatic view of an example of a rocket engineto which the invention is applicable.

DETAILED DESCRIPTION OF PREFERRED EMBODIMENTS

The description begins briefly with reference to FIG. 8 relating to anexample of an expander cycle rocket engine to which the invention isparticularly applicable.

A stream of fuel, such as hydrogen, stored in a tank 80 is pumped fromthe tank via a pipe 82 and a valve 81 by means of a pump 83 and itpasses along a pipe 84 to the wall 47 of the nozzle 40 where it flowsalong the nozzle in a network of tubes constituting a regeneratorcircuit 85 that lowers the temperature of the walls 47 of the topportion of the divergent section 43, of the nozzle throat 42, and of thecombustion chamber 41 by heating the fuel propellant that is extractedfrom the regenerator circuit 85 via a pipe 86. Turbines 87 and 77,forming portions of turbopumps for pumping the fuel propellant by meansof the pump 83 and for pumping the oxidizer propellant by means of thepump 73, receive propellant from the pipes 86, 97, and 98, whichpropellant has been gasified by the regenerator circuit 85 and has beenexhausted at the outlet from the turbines 87 and 77 via a duct 99leading to the combustion chamber 41. The entire flow of fuel propellantflowing in the pipe 86 is injected by the injector 45 into thecombustion chamber 41. The oxidizer propellant taken from the propellanttank 70 is pumped via a pipe 72 and a valve 71 using a pump 73, and itpasses via the pipe 74 to an injector 44 in order to be injected intothe combustion chamber 41.

In conventional manner, the regenerator circuit 85 may be constituted bya network of metal tubes fitted onto a shell constituting the wall ofthe combustion chamber and all or part of the rocket engine nozzle.

The invention is not limited to nozzles that incorporate the regeneratorcircuit 85, which is given merely by way of example.

The invention applies to any type of divergent section 43, and may beapplied, for example, to a rocket engine having a sheet of juxtaposedcooling tubes that are welded or brazed together, and that directlyconstitute at least a portion of the divergent nozzle. Nevertheless, theinvention is preferably applied to divergent nozzles made ofthermostructural composite material, e.g. of the carbon-carbon type,carbon-ceramic type, or ceramic-ceramic type, which nozzles mayoptionally be expandable.

The invention may also apply to a rocket engine operating with someother type of cooling circuit, e.g. an auxiliary hydraulic circuitdistinct from the main propellant streams being fed to the engine.

With a divergent section 43 that is not cooled and that is made ofcomposite material, when the rocket engine is in operation, it emitshigh levels of thermal radiation outwards. The power emitted by thedivergent section 43 is generally greater than 250 kW/m² of thedivergent section.

In accordance with the invention, a heater device 120 including at leastone duct conveying an inert fluid is arranged outside the nozzle 40 inthe immediate proximity thereof, but without making contact, in order torecover the energy of the thermal radiation that is emitted when therocket engine is in operation and thereby heat the inert fluid, whichmay for example be helium.

In FIG. 8, there can be seen a tank of helium 60 that is connected by apipe 62 and a valve 61 to the heater device 120, and a pipe 63 thatconnects the heater device to the liquid oxygen tank 70 in order toprovide the tank 70 with helium for pressurizing it.

FIG. 8 shows a heater device 120 of toroidal shape, but variousembodiments are possible for this radiant heat exchanger that serves torecover the radiant energy from the rocket engine and also to pressurizeone or more propellant tanks while obtaining an improvement of severaltens of kilograms in terms of payload in comparison with conventionalhelium pressurization systems that do not use a radiant heat exchanger.

The heater recovering the radiant energy coming from a rocket engine ismade up of a metal structure in which the fluid for heating flows. Theheater is dimensioned as a function of the available radiant power thatit is capable of absorbing in its location. For constant geometry, itsabsorption power depends both on its absorption properties and on anoverall visibility factor between the heater and the rocket engine. Thatis why a fine layer that is strongly absorbent from a thermal radiationpoint of view is deposited on the facing walls of the source of radiantpower. The strong absorption of the deposit makes it possible to avoidreflecting too much of the incident flux coming from the rocket engineso as to avoid creating any points of overheating. The fluid flowing inthe heat exchanger heats by conventional convection, thereby alsoserving to stabilize the temperature of the walls of the heater.

The environment of the heater is made up of:

-   -   the vacuum of space (at a temperature of 3 K); and    -   the rocket engine, itself comprising at least one part of        axisymmetrical shape that is hot in operation (T>1000 K).

In one possible embodiment, the heater device 20 comprises a metal platein the form of a frustoconical sector that extends around the divergentnozzle (represented diagrammatically by a truncated cone 30 in FIGS. 1and 3) over an angle σ lying in the range 30° to 360°.

As can be seen in FIG. 2, the plate 21 of the heater 20 includes anetwork of continuous channels 22 of rectangular or circular section. Afine layer 23 of strongly absorbent material from the thermal radiationpoint of view is formed on the wall of the plate 21 that faces theradiant power source 30.

The plate 21 of the heater 20 may present thickness lying in the rangeapproximately 5 millimeters (mm) to 15 mm.

In another embodiment, the heater 120 comprises a torus surrounding thenozzle as represented diagrammatically in FIG. 4 by a truncated cone 30.

The toroidal heater 120 may be placed around the divergent section ofthe nozzle, the throat of the nozzle, or the wall of the combustionchamber. A plurality of toruses sharing the axis of the nozzle may bejuxtaposed in stages along the radiant power source 30.

As in the embodiment of FIGS. 1 to 3, the torus(es) 120 possess(es) anabsorbent deposit at least on the outside faces of their walls wherethey face the radiant power source 30.

In another embodiment, as shown in FIG. 5, the heater 220 comprises atube that is wound helically around the radiant power source 30 asconstituted by the divergent section of the nozzle, and representeddiagrammatically by a truncated cone.

Under such circumstances, and as shown in FIG. 6, it is advantageous toimplement a heater device 320 that comprises a plurality of tubes, e.g.four tubes 321, 322, 323, and 324, that are wound helically around thedivergent section of the nozzle that is likewise representeddiagrammatically in FIG. 6 by a truncated cone.

In the embodiments of FIGS. 5 and 6, the helical tubes 220, 321, 322,323, and 324 all possess a layer of radiation-absorbent material ontheir outside faces, at least on the part facing towards the radiantpower source 30.

With the helical tubes 321 to 324 of FIG. 6, it is possible optionallyto make provision for some of the tubes to carry a counterflow, in orderto vary the efficiency of the heater 320. The helical shape of the tubessurrounding the axis of the hot source 30 makes it possible to make themmore uniform in temperature in order to avoid any overheating.

At equivalent fluid section and equivalent flow rate, the heat exchangecoefficient between the fluid traveling along the tubes 321 to 324 andthe walls of the tubes is increased and the heat exchanger section isgreater. The efficiency of this type of heater is thus much greater thanin a single torus heater as shown in FIG. 4, and its operating range isextended.

FIG. 7 shows a variant embodiment in which a heater device 420 comprisesfour twisted-together tubes 421 to 424 wound about a circular axis, i.e.included in the geometrical envelope of a torus, but that could equallywell be included in a helix having the same axis as the axis of thedivergent nozzle 30.

The number of tubes that are spiral-wound around a common circular axismay be other than four. It is thus possible to have pairs, triads,quads, or some greater number of tubes interleaved amongst one another.

It is also possible to have a plurality of different assemblies oftoroidal tubes 120 or of groups of tubes 420 staged along the entireheight of the divergent nozzle.

The invention claimed is:
 1. A cryogenic-propellant rocket enginecomprising: at least a first tank for a first liquid propellant; asecond tank for a second liquid propellant; a third tank for an inertfluid that is used to pressurize at least one of the first and secondtanks; an axisymmetrical nozzle including a combustion chamber, a devicefor injecting first and second liquid propellants into the combustionchamber, a nozzle throat, and a divergent section; a heater deviceincluding at least one duct for conveying the inert fluid and arrangedoutside the nozzle in immediate proximity of the nozzle, but withoutmaking contact with the nozzle, exclusively around said divergentsection, to recover energy of thermal radiation emitted when the rocketengine is in operation and to heat the inert fluid; and wherein theheater device comprises a metal structure in which the inert fluid forheating flows and a fine layer that is strongly absorbent from a thermalradiation point of view deposited at least on walls that face the nozzleconstituting a radiant power source.
 2. A rocket engine according toclaim 1, wherein the heater device comprises at least one torussurrounding the divergent section of the nozzle.
 3. A rocket engineaccording to claim 2, wherein the heater device comprises two to fourtubes surrounding the divergent section of the nozzle.
 4. A rocketengine according to claim 3, wherein the heater device comprises fourtwisted-together tubes wound around a circle having an axis thatcoincides with an axis of the divergent section.
 5. A rocket engineaccording to claim 3, wherein the tubes are arranged for providing acounterflow of the inert fluid in at least one tube.
 6. A rocket engineaccording to claim 1, wherein the heater device comprises at least onetube wound helically around the divergent section of the nozzle.
 7. Arocket engine according to claim 6, wherein the heater device comprisestwo to four tubes surrounding the divergent section of the nozzle.
 8. Arocket engine according to claim 7, wherein the heater device comprisesfour twisted-together tubes wound around a helix having an axis thatcoincides with an axis of the divergent section.
 9. A rocket engineaccording to claim 7, wherein the tubes are arranged for providing acounterflow of the inert fluid in at least one tube.
 10. A rocket engineaccording to claim 1, wherein the inert fluid is helium.
 11. A rocketengine according to claim 10, further comprising pipes making aconnection between the heater device and a liquid oxygen tankconstituting said at least a first tank for a first liquid propellant tosupply it with pressurizing helium.
 12. A rocket engine according toclaim 1, wherein said at least a first tank for a first liquidpropellant includes at least one tank of liquid oxygen.
 13. A rocketengine according to claim 12, further comprising pipes making aconnection between the heater device and the liquid oxygen tank tosupply it with pressurizing helium.
 14. A rocket engine according toclaim 1, wherein the divergent section of the nozzle is made ofthermostructural carbon-carbon, carbon-ceramic, or ceramic-ceramiccomposite material.